Control method for controlling a yaw angle and a roll angle of a vertical take-off aircraft

ABSTRACT

A control method for controlling a yaw angle γ z  and a roll angle γ x  of a vertical take-off aircraft comprising at least two drive groups arranged in opposite side regions of the aircraft so as to be spaced apart from a fuselage of the aircraft is presented. Each drive group comprises at least one first drive unit. The first drive unit is arranged so as to be spaced apart from the fuselage to pivot about a pivot angle α into a horizontal flight position and a vertical flight position.

TECHNICAL FIELD

The disclosure relates to a control method for controlling a yaw angleand a roll angle of a vertical take-off aircraft comprising at least twodrive groups arranged in opposite side regions of the aircraft so as tobe spaced apart from a fuselage of the aircraft, each drive groupcomprising at least one first drive unit, the first drive unit beingarranged so as to be spaced apart from the fuselage to pivot about apivot angle α into a horizontal flight position and a vertical flightposition.

BACKGROUND

Vertical take-off aircraft are used, inter alia, as drones and in themilitary field. These aircraft usually comprise two wings arranged onopposite sides of a fuselage, with two drive units being arranged so asto be pivotally mounted on each of the wings in supporting elements,such as gondolas, which are rigidly connected to the wings and areadapted to the intended purpose. Aircraft are also known in which thereis no separate fuselage and the wings are formed by two half-wings thatare symmetrical along the longitudinal axis, with two drive units beingarranged so as to be pivotally mounted on each of the half-wings insupporting elements which are rigidly connected to the half-wings andare adapted to the intended purpose.

In addition, vertical take-off aircraft are known from the prior art inwhich the drive units are pivotally mounted directly on the wings, forexample on a supporting structure extending within the wing. A verticaltake-off aircraft of this type is described in WO 2014/016226 A1. Inthis vertical take-off aircraft, it is provided that, in the horizontalflight position, the first drive unit is arranged on the wing above thewing surface and the second drive unit is arranged on the wing below thewing surface, and that, in the vertical flight position, the first driveunit and the second drive unit are arranged in an approximatelyhorizontal plane. In this way, in the vertical flight phase close to theground, a uniform ground effect of the first and second drive unit isachieved, such that smoother flight behavior is achieved, in particularin the take-off and landing phase. In the horizontal flight position,the first drive unit and the second drive unit do not flow against oneanother, and therefore this does not bring about any loss of efficiency.

In order to control the yaw angle and roll angle, in particular ofmulticopters, control methods are known in which the desired yaw androll is achieved by suitably actuating the drive units, which are notpivotable on multicopters. In this case, the power provided by therespective drive units is predetermined individually to each drive, inorder to generate the yaw and roll by means of the differences in liftand torque that are generated in this way.

An orientation of the aircraft in three-dimensional space is usuallydescribed by the roll, pitch and yaw angles. In this case, the differentangles describe rotational angles of the aircraft starting from a zeroposition which may, for example, correspond to the orientation of theaircraft when standing on the ground about a longitudinal, transverseand vertical axis of the aircraft.

SUMMARY

The problem addressed by the invention is considered that of providing acorresponding control method for controlling the yaw angle and rollangle by suitably actuating the drive units for vertical take-offaircraft, by means of which control is possible both in vertical flightas well as during the transition into horizontal flight, and inhorizontal flight.

This problem is solved by a control method for controlling a yaw angleand a roll angle of a vertical take-off aircraft of the type describedat the outset,

-   -   wherein power generated by each of the drive units is adapted in        order to reach a predetermined target yaw angle and a        predetermined target roll angle,    -   wherein, in a determination step, a first yaw control parameter        g₁ and a second yaw control parameter g₂ as well as a first roll        control parameter r₁ and a second roll control parameter r₂ are        determined, wherein the first yaw control parameter and the        first roll control parameter are vertical control parameters for        reaching the target yaw angle and the target roll angle in the        vertical flight position, wherein the second yaw control        parameter and the second roll control parameter are horizontal        control parameters for reaching the target yaw angle and the        target roll angle in the horizontal flight position,    -   wherein, in a subsequent superimposing step, an actuation        parameter is determined, by means of a superimposing rule for        each drive unit, from the vertical control parameters and the        horizontal control parameters on the basis of the pivot angle,    -   and wherein a power of the drive units is then predetermined        taking into account the actuation parameters.

Different actuation parameters are required in all the pivot positionsof the drive units in order to achieve the desired yaw and rollbehavior. By means of the continuous calculation of the horizontalcontrol parameters and the vertical control parameters as well assuperimposition of the parameters determined for the horizontal flightand vertical flight, the control method can be implemented particularlysimply, since the control method only has to be designed for the twoextreme horizontal and vertical flight positions. In order to determinethe horizontal control parameters and the vertical control parameters,control and regulation parameters that are known from the prior art canbe used. The actuation parameters may for example be power differencevalues that describe the deviation from the total power required by thedrive units to reach the target yaw and roll angles, the total powercorresponding to the power requirement requested by a pilot. Theactuation parameters are advantageously determined from roll controlparameters and yaw control parameters which represent absolute valuesfor the required power difference values in horizontal flight andvertical flight.

It is advantageously provided that each drive group comprises a firstdrive unit and a second drive unit, the first drive unit and the seconddrive unit each being arranged so as to be spaced apart from thefuselage to pivot about a pivot angle α into a horizontal flightposition and a vertical flight position. It is advantageously providedthat the first drive units and/or the second drive units are arranged onopposite wings of the aircraft. It is, however, also possible that thedrive units are arranged on the fuselage, advantageously by means ofsupporting frames, so as to be spaced apart from the fuselage. The driveunits may advantageously comprise rotors.

In a simplification of the described method, it is advantageouslyprovided that the second yaw control parameter is determined on thebasis of the first yaw control parameter by multiplication by a yawfactor and/or in that the second roll control parameter is determined onthe basis of the first roll control parameter by multiplication by aroll factor. It is also possible that the respective first controlparameters are determined by multiplying the respective second controlparameters by the associated factor.

It is advantageously provided that, in the determining step, an actualyaw angle and an actual roll angle are determined, and that the controlparameters are each determined by means of a control algorithm startingfrom the target yaw angle and the target roll angle as well as theactual yaw angle and the actual roll angle. The actual yaw angle and theactual roll angle may advantageously be detected by means of suitablesensors and may be transferred to a microcontroller or flight controllerexecuting the control algorithm and/or the control method.

It is advantageously provided that the first yaw control parameter isdetermined on the basis of the target yaw angle and the actual yaw angleusing a first yaw control algorithm and/or that the second yaw controlparameter is determined on the basis of the target yaw angle and theactual yaw angle using a second yaw control algorithm and/or that thefirst roll control parameter is determined on the basis of the targetroll angle and the actual roll angle using a first roll controlalgorithm and/or that the second roll control parameter is determined onthe basis of the target roll angle and the actual roll angle using asecond roll control algorithm. By using a plurality of separate controlalgorithms to determine the control parameters, the design of theindividual control algorithms can be significantly simplified, sinceonly SISO systems have to be taken into account. Coupling between theoutput variables can then optionally be taken into account by theadvantageously non-linear superimposition of the different controlparameters.

In a particularly advantageous embodiment of the method, it is providedthat the first yaw control algorithm and/or the second yaw controlalgorithm and/or the first roll control algorithm and/or the second rollcontrol algorithm is a linear controller having a P or PD proportion.The use of linear controllers having a P or PD proportion isparticularly simple. Advantageously, the control algorithms mayadditionally also have an I proportion.

Advantageously, the first yaw control parameter g₁ is determined using aPD controller starting from the actual yaw angle y_(γ) _(z)advantageously determined using suitable sensors and from the actual yawrate {dot over (y)}_(γ) _(z) advantageously also detected using sensorsand starting from the predetermined target yaw angle w_(γ) _(z) , inaccordance with the following provision:

g ₁=(P _(g) ₁ ·(w _(γ) _(z) −y _(γ) _(z) )−{dot over (y)} _(γ) _(z) )·D_(g) ₁

The factor P_(g) ₁ represents the P proportion of the PD controller andthe factor D_(g) ₁ represents the D proportion of the PD controller fordetermining the first yaw control parameter g₁. In a comparable manner,the first roll control parameter r₁, the second yaw control parameter g₂and the second roll control parameter r₂ are determined in accordancewith the following provisions, with the actual roll angle y_(γ) _(x)advantageously detected using suitable sensors, the actual roll rate{dot over (y)}_(γ) _(x) likewise advantageously detected by sensors andthe predetermined target roll angle w_(γ) _(x) additionally being used:

r ₁=(P _(r) ₁ ·(w _(γ) _(x) −y _(ν) _(x) )−{dot over (y)} _(γ) _(x) )·D_(r) ₁

g ₂=(P _(g) ₂ ·(w _(γ) _(z) −y _(γ) _(z) )−{dot over (y)} _(γ) _(z) )·D_(g) ₂

r ₂=(P _(r) ₂ ·(w _(γ) _(x) −y _(ν) _(x) )−{dot over (y)} _(γ) _(x) )·D_(r) ₂

It is advantageously provided that the P proportion P_(g) ₁ and the Dproportion D_(g) ₁ for determining the first yaw control parameter g₁correspond to the P proportion P_(r) ₂ and the D proportion D_(r) ₂ ofthe PD controller for determining the second roll control parameter r₂,and that the P proportion P_(r) ₁ and the D proportion D_(r) ₁ fordetermining the first roll control parameter r₁ correspond to the Pproportion P_(g) ₂ and the D proportion D_(g) ₂ of the PD controller fordetermining the second yaw control parameter g₂.

In order to determine the actual roll angle y_(γ) _(x) , the actual rollrate {dot over (y)}_(γ) _(x) , the actual yaw angle y_(γ) _(z) and theactual yaw rate {dot over (y)}_(γ) _(z) , the aircraft advantageouslycomprises a gyroscope, an acceleration sensor and a compass, with therequired angles and rates being determined on the basis of methods knownfrom the prior art, such as a Kalman filter composed of measuredvariables detected by these sensors.

It is also provided that the horizontal control parameters and thevertical control parameters are continuously each determined using acontrol algorithm or a common control algorithm. Each control algorithmused may be a linear or non-linear controller.

It is advantageously provided that, in the superimposing step, thevertical control parameters and the horizontal control parameters areeach multiplied by a drive-unit-specific and pivot-angle-specificevaluation function and the actuation parameters for each drive unit aredetermined by a linear combination of the vertical control parametersmultiplied by the drive-unit-specific and pivot-angle-specificevaluation function and the horizontal control parameters multiplied bythe drive-unit-specific and pivot-angle-specific evaluation function.Advantageously, the evaluation function is a non-linear function basedon the pivot angle. By means of the non-linear evaluation and subsequentlinear combination, unconsidered couplings between the control loop canalso advantageously be considered, in particular when using a pluralityof separate controllers for determining the first and second yaw androll control parameters.

It is advantageously provided that the evaluation function of thevertical control parameters is the cosine of the pivot angle and thatthe evaluation function of the horizontal control parameters is the sineof the pivot angle. It has been shown that the superimposition usingsine and cosine functions can achieve particularly stable flightbehavior, in particular during the transition from the vertical flightposition into the horizontal flight position, and vice versa.

In a particularly advantageous embodiment of the method, it is providedthat a yaw angle and a roll angle are defined in a clockwise mannerabout a vertical axis and a longitudinal axis, respectively, of theaircraft, in the superimposing step, the actuation parameter AP₁ of afirst drive unit arranged to the left of the longitudinal axis in a planview of the aircraft being calculated according to the following model:

AP ₁=cos(α)·r ₁+cos(α)·g ₁−sin(α)·r ₂+sin(α)·g ₂,  (1)

in the superimposing step, the actuation parameter AP₂ of a first driveunit arranged to the right of the longitudinal axis in a plan view ofthe aircraft being calculated according to the following model:

AP ₂=−cos(α)·r ₁−cos(α)·g ₁+sin(α)·r ₂−sin(α)·g ₂,  (2)

in the superimposing step, the actuation parameter AP₃ of a second driveunit arranged to the left of the longitudinal axis in a plan view of theaircraft being calculated according to the following model:

AP ₃=cos(α)·r ₁−cos(α)·g ₁+sin(α)·r ₂+sin(α)·g ₂,  (3)

in the superimposing step, the actuation parameter AP₄ of a second driveunit arranged to the right of the longitudinal axis in a plan view ofthe aircraft being calculated according to the following model:

AP ₄=−cos(α)·r ₁+cos(α)·g ₁−sin(α)·r ₂−sin(α)·g ₂.  (4)

If the drive units comprise rotors, a rotational direction of the rotorsis advantageously taken into account when determining the actuationparameters.

In order to determine power actuation values, by means of which theindividual drive units are subsequently actuated, starting from the thusdetermined actuation parameters, which advantageously represent powerdifference values, it is provided that power actuation values u₁, u₂,u₃, u₄ of the drive units, by means of which the drive units areactuated, are calculated as follows, taking into account a powerrequirement variable F and a pitch parameter n, in order to generate thedesired power of the individual drive units:

u ₁ =F−n+AP ₁,  (5)

u ₂ =F−n+AP ₂,  (6)

u ₃ =F+n+AP ₃,  (7)

u ₄ =F+n+AP ₄.  (8)

The power requirement variable may for example be the total powerrequested by a pilot. The pitch parameter advantageously describes apower difference value which is required to reach a predetermined targetpitch angle.

It is advantageously provided that, in the horizontal flight position,the first drive units are arranged in the direction of the vertical axisso as to be spaced apart from the second drive units, and that, in thevertical flight position, the first drive units are arranged in thedirection of the longitudinal axis so as to be spaced apart from thesecond drive units. It is advantageously provided that, in thehorizontal flight position, the first drive units are arranged above anupper wing surface and the second drive units are arranged below a lowerwing surface, and that, in the vertical flight position, the first driveunits and the second drive units are arranged in front of and behind thewing in the horizontal flight direction.

Other advantageous configurations of the method are explained in greaterdetail with reference to an embodiment shown in the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The FIGURE schematically shows a vertical take-off aircraft.

DETAILED DESCRIPTION

The FIGURE schematically shows a vertical take-off aircraft 1. Theaircraft 1 comprises two drive groups 3 arranged on opposite wings 2 ofthe aircraft 1, each drive group 3 comprising a first drive unit 4, 5and a second drive unit 6, 7. The first drive unit 4, 5 and the seconddrive unit 6, 7 are each arranged on the wing 2 so as to pivot about apivot angle α into a horizontal flight position and a vertical flightposition. In the FIGURE, the drive units 4, 5, 6, 7 are in thehorizontal flight position. The first drive units 4, 5 are arranged onthe wings 2 above an upper wing surface 8 and the second drive units 6,7 are arranged on said wings below a lower wing surface 9. When thedrive units 4, 5, 6, 7 are pivoted into the vertical flight position,the first drive units 4, 5 and the second drive units 6, 7 are arrangedin front of and behind the wings 2 in the horizontal flight direction. Ayaw angle γ_(z), a roll angle γ_(x) and a pitch angle γ_(y) are definedin a clockwise manner about a vertical axis, a longitudinal axis and atransverse axis, respectively, of the aircraft 1.

The drive units 4, 5, 6, 7 each comprise rotors. The rotors of the firstdrive unit 4 and the second drive unit 7 rotate anti-clockwise and therotors of the first drive unit 5 and the second drive unit 6 rotateclockwise.

In order to control the yaw angle γ_(z) and the roll angle γ_(x), in adetermining step of a control method, first and second yaw and rollcontrol parameters g₁, g₂, r₁, r₂ are first determined starting from apredetermined target yaw angle w_(y), and a predetermined target rollangle w_(γ) _(x) by means of linear controllers PD₁, PD₂, PD₃, and PD₄.Actuation parameters u₁, u₂, u₃, u₄ are then determined for each driveunit 4, 5, 6, 7 in a superimposing step from the roll control parametersg₁, g₂, r₁, r₂. In the FIGURE, the determination of the actuationparameter u₁ for the first drive unit 4 is shown by way of example. Thedetermination takes place on the basis of the above-described formulas 1to 8.

1.-10. (canceled)
 11. A control method for controlling a yaw angle(γ_(z)) and a roll angle (γ_(x)) of a vertical take-off aircraft (1)comprising at least two drive groups (3) arranged in opposite sideregions of the aircraft (1) so as to be spaced apart from a fuselage ofthe aircraft, wherein each drive group (3) comprises at least one firstdrive unit (4, 5), wherein the first drive unit (4, 5) is arranged so asto be spaced apart from the fuselage to pivot about a pivot angle (α)into a horizontal flight position and a vertical flight position,wherein power generated by each of the drive units (4, 5, 6, 7) isadapted in order to reach a predetermined target yaw angle (w_(γ) _(z) )and a predetermined target roll angle (w_(γ) _(x) ), wherein, in adetermination step, a first yaw control parameter (g₁) and a second yawcontrol parameter (g₂) as well as a first roll control parameter (r₁)and a second roll control parameter (r₂) are determined, wherein thefirst yaw control parameter (g₁) and the first roll control parameter(r₁) are vertical control parameters for reaching the target yaw angle(w_(γ) _(z) ) and the target roll angle (w_(γ) _(x) ) in the verticalflight position, wherein the second yaw control parameter (g₂) and thesecond roll control parameter (r₂) are horizontal control parameters forreaching the target yaw angle (w_(γ) _(z) ) and the target roll angle(w_(γ) _(x) ) in the horizontal flight position, wherein, in asubsequent superimposing step, an actuation parameter is determined, bya superimposing rule for each drive unit (4, 5, 6, 7), from the verticalcontrol parameters and the horizontal control parameters on the basis ofthe pivot angle (α), and wherein a power of the drive units (4, 5, 6, 7)is then predetermined taking into account the actuation parameters. 12.The control method according to claim 11, wherein each drive group (3)comprises a first drive unit (4, 5) and a second drive unit (6, 7),wherein the first drive unit (4, 5) and the second drive unit (6, 7)each are arranged so as to be spaced apart from the fuselage to pivotabout the pivot angle (α) into a horizontal flight position and avertical flight position.
 13. The control method according to claim 11,wherein the second yaw control parameter (g₂) is determined on the basisof the first yaw control parameter (g₁) by multiplication by a yawfactor and/or wherein the second roll control parameter (r₂) isdetermined on the basis of the first roll control parameter (r₁) bymultiplication by a roll factor.
 14. The control method according toclaim 11, wherein, in the determining step, an actual yaw angle (y_(γ)_(z) ) and an actual roll angle (y_(γ) _(x) ) are determined, andwherein the control parameters are each determined by a controlalgorithm starting from the target yaw angle (w_(γ) _(z) ) and thetarget roll angle (w_(γ) _(x) ) as well as the actual yaw angle (y_(γ)_(z) ) and the actual roll angle (y_(γ) _(x) ).
 15. The control methodaccording to claim 14, wherein the first yaw control parameter (g₁) isdetermined on the basis of the target yaw angle (w_(γ) _(z) ) and theactual yaw angle (y_(γ) _(z) ) using a first yaw control algorithm (PD₂)and/or wherein the second yaw control parameter (g₂) is determined onthe basis of the target yaw angle (w_(γ) _(z) ) and the actual yaw angle(y_(γ) _(z) ) using a second yaw control algorithm (PD₄) and/or whereinthe first roll control parameter (r₁) is determined on the basis of thetarget roll angle (w_(γx)) and the actual roll angle (y_(γx)) using afirst roll control algorithm (PD₁) and/or wherein the second rollcontrol parameter (r₂) is determined on the basis of the target rollangle (w_(γ) _(x) ) and the actual roll angle (y_(γ) _(x) ) using asecond roll control algorithm (PD₃).
 16. The control method according toclaim 15, wherein the first yaw control algorithm (PD₂) and/or thesecond yaw control algorithm (PD₄) and/or the first roll controlalgorithm (PD₁) and/or the second roll control algorithm (PD₃) is alinear controller having a P or PD proportion.
 17. The control methodaccording to claim 11, wherein, in the superimposing step, the verticalcontrol parameters and the horizontal control parameters are eachmultiplied by a drive-unit-specific and pivot-angle-specific evaluationfunction and the actuation parameters for each drive unit (4, 5, 6, 7)are determined by a linear combination of the vertical controlparameters multiplied by the drive-unit-specific andpivot-angle-specific evaluation function and the horizontal controlparameters multiplied by the drive-unit-specific andpivot-angle-specific evaluation function.
 18. The control methodaccording to claim 17, wherein the evaluation function of the verticalcontrol parameters is the cosine of the pivot angle (α) and wherein theevaluation function of the horizontal control parameters is the sine ofthe pivot angle (α).
 19. The control method according to claim 18,wherein the yaw angle (γ_(z)) and the roll angle (γ_(x)) are defined ina clockwise manner about a vertical axis and a longitudinal axis,respectively, of the aircraft (1), wherein, in the superimposing step, afirst actuation parameter (AP₁) of a first drive unit (4) arranged tothe left of the longitudinal axis in a plan view of the aircraft iscalculated according to the following model:AP ₁=cos(α)·r ₁+cos(α)·g ₁−sin(α)·r ₂+sin(α)·g ₂, wherein, in thesuperimposing step, a second actuation parameter (AP₂) of a first driveunit (5) arranged to the right of the longitudinal axis in a plan viewof the aircraft is calculated according to the following model:AP ₂=−cos(α)·r ₁−cos(α)·g ₁+sin(α)·r ₂−sin(α)·g ₂, wherein, in thesuperimposing step, a third actuation parameter (AP₃) of a second driveunit (6) arranged to the left of the longitudinal axis in a plan view ofthe aircraft is calculated according to the following model:AP ₃=cos(α)·r ₁−cos(α)·g ₁+sin(α)·r ₂+sin(α)·g ₂, wherein, in thesuperimposing step, a fourth actuation parameter (AP₄) of a second driveunit (7) arranged to the right of the longitudinal axis in a plan viewof the aircraft is calculated according to the following model:AP ₄=−cos(α)·r ₁+cos(α)·g ₁−sin(α)·r ₂−sin(α)·g ₂.
 20. The controlmethod according to claim 19, wherein power actuation values (u₁, u₂,u₃, u₄) of the drive units (4, 5, 6, 7), by means of which the driveunits (4, 5, 6, 7) are actuated, are calculated as follows, taking intoaccount a power requirement variable (F) and a pitch parameter (n), inorder to generate a desired power of the individual drive units (4, 5,6, 7):u ₁ =F−n+AP ₁,u ₂ =F−n+AP ₂,u ₃ =F+n+AP ₃,u ₄ =F+n+AP ₄.